In the present era of environmental awareness, the gas turbine engine designer, and particularly the designer of such engines for aircraft propulsion, is faced with the dilemma of reducing engine pollutants while sacrificing the minimum engine performance. One type of pollution which recently has received considerable attention is noise.
Gas turbine engine noise is generated from two primary sources. First, there is that associated with the viscous shearing of rapidly moving gases exhausted into the relatively quiescent surrounding atmosphere. In turbofan aircraft engines, such gases are emitted from the fan and core nozzles at the rear of the engine. Various approaches have been utilized to reduce this "shear" noise, most approaches incorporating mixers to comingle fan and exhaust gases with each other and with the surrounding environment.
The second source of noise, and the one to which the present invention is directed, is generated by the rotating turbomachinery itself. This results from the relative motion between the rapidly rotating blade rows and the interflowing gas stream. The noise is affected by such parameters as blade rotational speed, blade-to-blade spacing, blade geometry, and by the proximity of stationary hardware to such rotating blade rows, as in the case of an outlet guide vane arrangement. Another example of the latter condition occurs in typical multistage axial compressors where stationary blade rows are alternated with rotating blade rows. Some of the noise generated in this manner can be absorbed and suppressed by means of acoustic or sound absorbing paneling disposed about the periphery of the nacelle enclosing the rotating turbomachinery. Such sound absorbing material is well known in the art. However, because of the close proximity of the fan or compressor to the inlet frontal plane, and the lack of acoustic shielding in the forward direction, a significant percentage of noise propagates forward from the gas turbine inlet duct.
Prior attempts to solve this problem have concentrated on the application of sound absorbing material to the inlet duct inner wall. This does little to attenuate unreflected noise propagating in the axially forward direction. Additional benefits have been obtained by providing coaxial, circumferential rings of sound absorbent material within the inlet. However, such rings produce a loss of inlet total pressure and, therefore, bring about performance losses which remain throughout the engine operating envelope even when noise propagation presents no hazard or nuisance to inhabitants below.
Another concept incorporates an axially translating wedge-shaped scoop on the bottom of the inlet duct to selectively reduce the downward transmission of noise from the inlet. However, this configuration is inadequate for two reasons. First, it has been shown that an inlet incorporating such a scoop has a poor pressure recovery characteristic (i.e., it is inherently a high loss system). Secondly, and somewhat related to the foregoing problem is that the total pressure pattern is highly distorted, as for example in the plane of a gas turbine fan stage disposed within the duct. While the former characteristic results in degraded engine performance, the latter may, under certain conditions, cause excessive fan blade stresses and possible destruction of the rotating turbomachinery.
Yet another approach has been to extend axially forward the lower cylindrical half of the inlet duct. In side profile, this results in a stepped duct wall contour. Although the configuration tends to reduce noise level, it is aerodynamically undesirable from the inlet recovery and distortion aspects discussed hereinbefore.
The problem facing the gas turbine designer is, therefore, to provide a means for attenuating noise emanating from the duct without incurring overall performance penalties.